1. Field of the Invention
This invention relates generally to a spacecraft system for converting solar array power into thruster power and, more particularly, to a spacecraft system that regulates the maximum power point output of a solar array so as to convert the maximum solar array power to a maximum usable current for a series of arcjet thrusters that maneuver the spacecraft.
2. Discussion of the Related Art
As is well understood, spacecrafts include a number of systems and loads that require a significant amount of power to be operated. For current spacecraft mission designs, multiple kilowatts of power is necessary to satisfy the requirements of the systems and loads. For this reason, spacecrafts usually incorporate one or more solar arrays consisting of photovoltaic cells that receive sunlight, and convert the sunlight to usable electrical power so as to enable the spacecraft to generate the necessary power to operate the spacecraft systems. One of the main spacecraft loads that will use the converted sunlight energy is the thrusters that maneuver the spacecraft. Thrusters are necessary propulsion devices of satellites and spacecrafts to provide such functions as orbit transfer, orbit adjustment, momentum management and station keeping. Other spacecraft loads include communication systems, heating systems, sensing systems, operating systems, etc.
The photovoltaic cells of a spacecraft solar array significantly degrade over the life of a spacecraft mission as a result of being exposed to the space environment. Consequently, the power that the solar array is able to generate is continually reduced through time. Further, the output power of the photovoltaic cells varies greatly with different temperatures. It is generally desirable to operate the spacecraft thrusters at or near the maximum solar array power output so as to use the most power available from the solar array power supply, and thus maintain the maximum thrust available to maneuver the spacecraft. By efficiently utilizing the power output of the solar array, spacecraft efficiency is increased so that less power is required for specific propulsion functions, and solar array size can be minimized.
For at least the above reasons, spacecrafts generally include a control mechanism that can track the maximum power output point of the solar array as it degrades with mission life and varies with temperature changes so that the thrusters operate at the maximum solar array voltage, and thus achieve maximum thrust efficiency and minimize solar array power to meet mission requirements. Accordingly, spacecrafts usually include a maximum power point tracking device that regulates the maximum power output of the solar array so as to effectively convert this power output to electrical energy usable by the propulsion system and loads of the spacecraft.
More support for utilizing a solar array maximum power point regulator can be gleaned from an analysis of FIG. 1. FIG. 1 shows the current and voltage relationship of the output of a solar array associated with a spacecraft and a series of load line curves of a power processing unit (PPU) that conditions the power applied to an arcnet thruster of the spacecraft. An arcjet thruster is a plasma or an ion type thruster, well known to those skilled in the art, that is more particularly adapted to lengthy spacecraft missions. Because the arcjet thruster is an ion thruster, the PPU is required in order to convert the solar array power to an appropriate current useable by the arcjet thruster. The solar array output line is a constant power output of the solar array and each load line curve represents a constant power output of the thruster. Below a reference voltage V.sub.R, the power output of the thruster is substantially linear and the power to the thruster is not regulated. When the voltage applied to the thruster reaches V.sub.R, V.sub.R must be less than the available solar array current to obtain regulation. This condition is illustrated by the load line curve 1 where point A represents the only operating point of the system at the intersection of the solar array power output line and the load line curve 1.
As the PPU load power is increased, the PPU power is represented by the load line curve 2. The load line curve 2 has three possible operating points at the intersection with the solar array power output line. For the load line curve 2, the point C is an unstable operating point as a result of the negative resistance characteristics of the PPU and the high resistance of the solar array at that point. Although the point D is a stable operating point, it is on the non-regulation portion of the curve 2, and thus is a non-desirable operating point. The desirable operating point is point B for this power output so the power output can be regulated.
The load line curve 3 represents a higher PPU power curve than the load line curve 2. As is apparent, point E is the maximum power point of the solar array. Arcjet thrusters are unique in the fact that the current applied to the thruster needs to be limited to a certain maximum current level so as to prevent the thruster from collapsing and/or burning out. When an ion arcjet thruster collapses, it is necessary to restart the thruster from an initial start condition. Therefore, any increase in power from the load line curve 3 will cause the power output of the PPU to be reduced to the point F. In this condition, the system power demand needs to be reduced to the load line curve 1 to allow a desirable operating point on the solar array curve.
As is apparent from an analysis of FIG. 1 and the corresponding discussion, active electronic systems are needed to operate arcjet thrusters of a spacecraft at or near the solar array maximum power point. Exceeding the maximum power point may cause thruster shut down and require a new start up procedure. This condition results in loss of thrust and possible loss of spacecraft control. Therefore, without active electronic control systems, the thrusters must be operated at a safe margin away from the maximum power point of the solar array. This margin can be considerable since an arcjet thruster is a noisy load, and the solar array maximum power point varies with temperature and mission life.
U.S. Pat. No. 4,794,272 issued to Bavaro et al., herein incorporated by reference, discloses a number of maximum power point tracking schemes for use in tracking the maximum power output of the solar array of a spacecraft. In the background discussion of Bavaro et al., a number of prior art solar array power regulators are disclosed that take advantage of analog processing, digital processing, or a priori techniques with respect to a known reference voltage. The analog processing technique involves a trial and error method that is based on the output power of the solar array being a continuous function of voltage and current with a single peak power point. The digital processing technique utilizes analog sensors and digital computers for performing the same type of function as the analog processing system. The a priori technique utilizes tests of solar arrays representing a regulated solar array so as to predict the peak power point of the regulated solar array.
The main portion of the Bavaro et al. patent is concerned with a power regulator that adjusts the operating point of a solar array and a battery as a function of the sensed output current of the particular power source. The regulator determines whether the battery is properly charging. If the battery is in an undercharging condition, the solar array operating point is adjusted so as to minimize the undercharge current. If the battery is in a charging condition, the operating point of the solar array is again adjusted so as to continue battery charging according to a predetermined charging scheme. The Bavaro et al. system requires determining phase changes in the output power of the solar array. Further, the Bavaro et al. system can only operate at the maximum power point of the solar array, and thus is not adaptable for operating at a power level that can be manipulated as desired.
A system 10 of a basic prior art approach for tracking the maximum power point of a solar array that could incorporate the schemes mentioned above is shown by a schematic block diagram in FIG. 2. Solar array power from a solar array 12 is processed through a series component referred to as a maximum power point tracker (MPPT) 14. The MPPT 14 is a control device that measures either deviation of the solar array power from the maximum power point of the solar array or the derivative of solar array power with respect to voltage of the power output signal from the solar array 12. Once the MPPT 14 finds the maximum power point of the solar array 12, the MPPT 14 uses a dithering process to maintain the output power of the MPPT 14 at this value regardless of variations in the solar array 12. As the MPPT 14 searches for the maximum power point of the solar array 12, the MPPT 14 will measure the voltage output of the solar array 12 to search for the maximum voltage from the solar array 12 in a perturbing manner. To accomplish this, the MPPT 14 must also determine either the time or phaseshift of the search reversal, or produce an output proportional to a searched error signal, depending on the type of system being used. Some systems also require that the MPPT 14 match the impedance of the solar array 12 with the impedance of the loads.
The maximum power value from the solar array 12 is transferred from the MPPT 14 to a power switching and logic circuit 16. The power switching and logic circuit 16 is a switching circuit that transfers the power from the solar array 12 to the spacecraft loads. Particularly, power from the solar array 12 is transferred through the power switching and logic circuit 16 to thrusters 18 for maneuvering the spacecraft. A propulsion powered conditioning circuit 20 converts the power signal from the solar array 12 to a form useable by the thrusters 18, as is well understood in the art. Also, solar array power is transferred through the power switching and logic circuit 16 to a variable power load 22. The variable power load 22 represents those spacecraft loads that require a different voltage than is provided by the voltage output of the power switching and logic circuit 16. Additionally, power from the solar array 12 is also applied through the power switching and logic circuit 16 to other spacecraft loads 24 through a spacecraft power conditioning circuit 26.
As is well understood in the art, the spacecraft loads may require more power than is able to be produced by the photovoltaic cells of the solar array 12 even when the MPPT 14 is operating at the maximum power point. Such times may occur when the photovoltaic cells of the solar array 12 are being eclipsed by some object, as would occur when the spacecraft travelled behind a planet or the like. Additionally, the solar array 12 may be designed to not be able to provide the necessary power when all or most of the spacecraft loads are operating simultaneously. For this reason, a spacecraft will include a battery 28 to augment the power output of the solar array 12. As is apparent, the output of the battery 28 is applied to the power switching and logic circuit 16, and thus, the battery power is distributed to the spacecraft loads in the same manner that the power from the solar array 12 is distributed. At times when the solar array 12 is capable of providing the power necessary to operate the spacecraft loads, the solar array power is also available to charge the battery 28 through the power switching and logic circuit 16 so as to maintain the power level of the battery 28 at a maximum capacity for times when the solar array 12 is unable to provide the necessary power to the spacecraft loads.
One of the greatest concerns when designing the systems necessary to operate a spacecraft is to limit the spacecraft weight as much as possible. For every reduction of weight of the spacecraft operating systems, greater payload capabilities can be realized with the same fuel requirements so as to increase the effectiveness for a particular mission. For example, the MPPT 14 and the battery 28 weigh hundreds of pounds in a multikilowatt spacecraft power system. Additionally, these elements add significantly to the overall cost of the spacecraft system 10. For certain kinds of spacecraft missions, such as interplanetary travel, the solar array 12 is generally capable of providing the power necessary to operate the spacecraft loads at all times. Therefore, it is possible to either eliminate the battery or significantly reduce the size of the battery.
From the discussion above, it is clear that it would be highly advantageous, especially in long term space travel missions, to eliminate the weight and expense of a maximum power point tracking device and battery in known spacecraft power systems. It would also be advantageous to provide a system that effectively tracked the maximum power point of a solar array so as to provide maximum current to an arcjet thruster. It would further be advantageous to provide a system that was capable of operating at any power point along the power output of the solar array. It is therefore, an object of the present invention to provide these advantages.